In aircraft construction attempts are being made increasingly to use, as load-bearing components, components that are made entirely or partially from fibre-reinforced composite materials, for example carbon fibre-reinforced plastics (CFP). For example DE 10 2007 062 111 A1 describes a crossmember structure made of carbon fibre-reinforced plastics material, which is used to support the individual panels of an aircraft floor system for dividing a passenger cabin from a cargo compartment disposed underneath the passenger cabin. It is further known for example from DE 10 2004 001 078 A1 to provide aircraft fuselage segments with a skin of a sandwich structure and with reinforcing elements (for example frames, stringers) made of fibre-reinforced composite materials.
During the manufacture—represented in FIG. 1—of an aircraft structural component 10 in the form of an aircraft fuselage segment from a fibre-reinforced composite material, reinforcing fibre layers impregnated with synthetic resin are first brought into a desired shape of a surface portion 12, which forms an aircraft skin, and a reinforcing portion 14, which forms a frame or stringer. The surface portion 12 and the reinforcing portion 14 are then joined to one another “wet in wet”, i.e. without prior curing of the synthetic resin forming a matrix of the fibre-reinforced composite material, in an autoclave. During the treatment in the autoclave a steel mould 16, which receives the surface portion 12 and the reinforcing portion 14, guarantees the dimensional stability of the external contours of the aircraft structural component 10. Furthermore, for additionally stabilizing the shape of the reinforcing portion 14 a tube 20 is introduced into a cavity 18 that is delimited by a region 12a of the surface portion 12 and by the reinforcing portion 14. The tube 20 is loaded with an internal pressure and therefore exerts a corresponding internal pressure on the region 12a of the surface portion 12 and the reinforcing portion 14 that delimit the cavity 18.
Particularly during the processing of composite materials having a matrix of a thermoset plastics material there is always the risk that during the curing process in the autoclave, which once started is no longer reversible, pores or stresses may arise in the matrix of the composite material and necessitate a secondary finishing and/or repair of the aircraft structural component 10. Particularly problematical, here, are component faults such as pores in a region of the aircraft structural component 10 adjoining the cavity 18, because repair of these component faults entails removing and then replacing the damaged component region.